Blunted leading edge fan blade for noise reduction

ABSTRACT

A blade of a turbomachinery rotor stage includes a leading edge with a blunted contour over at least a portion of the radial length thereof. The blunting is accomplished in a variety of ways, such as by removing material from a portion of the leading edge, by blowing air into the leading edge or by positioning a mechanical device such as an inflatable boot around a portion of the leading edge. The aerodynamic or mechanical blunting can be utilized during selective portions of flight, such as take-off and landing. The blunting of leading edges of each of a plurality of blades in this way provides significant reduction in noise generated during rotation of the associated turbomachinery rotor state.

Evans et al.

[ BLUNTED LEADING EDGE FAN BLADE FOR NOISE REDUCTION [75] Inventors:Robert C. Evans; Donald K.

Dunbar, both of Cincinnati, Ohio [73] Assignee: General ElectricCompany,

Cincinnati, Ohio [22] Filed: Aug. 29, I973 [21] Appl. No.: 392,508

Related US. Application Data [63] Continuation-in-part of Ser. No.194,285, Nov. 1,

1971, abandoned.

[52] [1.5. CI 416/223; 416/228 [51] Int. Cl. FOId 5/14 [58] Field ofSearch 416/223, 228; 415/119 [56] References Cited UNITED STATES PATENTS3,347,520 10/1967 Owczarek 446/228 X 3,403,893 10/1968 Stoffer 416/228 41 May 13, 1975 3,542,486 ll/l97O Kercher 416/90 PrimaryExaminer-Everette A. Powell, Jr.

Attorney, Agent, or Firm-Derek P. Lawrence; Lee H. Sachs [57] ABSTRACT Ablade of a turbomachinery rotor stage includes a leading edge with ablunted contour over at least a portion of the radial length thereof.The blunting is accomplished in a variety of ways, such as by removingmaterial from a portion of the leading edge, by blowing air into theleading edge or by positioning a mechanical device such as an inflatableboot around a portion of the leading edge. The aerodynamic or mechanicalblunting can be utilized during selective portions of flight, such astake-off and landing. The blunting of leading edges of each of aplurality of blades in this way provides significant reduction in noisegenerated during rotation of the associated turbomachinery rotor state.

5 Claims, 7 Drawing Figures BLUNTED LEADING EDGE FAN BLADE FOR NOISEREDUCTION This application is a continuation-in-part of application Ser.No. 194,285, filed Nov. 1, l97l now abandoned.

BACKGROUND OF THE INVENTION This invention relates generally to gasturbine engine noise reduction devices and, more particularly, to suchan engine which includes turbomachinery rotor blades having leadingedges which are contoured or otherwise altered to reduce the noisegenerated by such blades.

A vast amount of effort has been expended in recent years by gas turbineengine manufacturers and aircraft manufacturers in efforts to reduce thenoise levels associated with aircraft gas turbine engines. This work hasincluded both effort toward suppressing noise generated by the engineswith suppression devices such as acoustic panels positioned within theinlets and exhaust passageways of the engine, and also has includedeffort toward reducing the noise generation per se by means such asproper blade-vane spacing, proper blade lean,

etc.

The advent of high bypass ratio, large diameter turbofan engines, whichpower the new wide body jets, has resulted in a generation of gasturbine engines in which the noise produced by the large diameter fangreatly exceeds and shadows the noise produced by a core engine whichdrives the fan. While the overall noise levels of these engines havebeen greatly reduced when compared to prior generation turbojets, anydesign innovation which would significantly reduce noise generation ofthe fan without significantly altering the basic structure of the enginewould greatly enhance the saleability of such an engine. This is truebecause the reduction in noise generation would permit either furtherreduction of the overall noise level of the engine or would permit theelimination of the added weight and cost associated with a noisereduction panel previously used to maintain the engine at a certainnoise level.

SUMMARY OF THE INVENTION It is an object of the present inventiontherefore to provide a turbomachinery rotor blade which reduces thenoise generated by a turbofan engine.

Briefly stated, the above and similarly related objects are accomplishedby providing a blade for a gas turbine engine rotor stage which bladeincludes a leading edge with a blunted contour over at least a portionof the radial length thereof. That is, the basic aerodynamic contour ofthe leading edge is made blunt or flatter than normal. The blunting maybe accomplished in a variety of ways, such as by removing material fromat least a portion of the leading edge, by blowing air into the leadingedge, or by positioning a mechanical device such as an inflatable bootaround a portion of the leading edge.

DESCRIPTION OF THE DRAWINGS While the specification concludes with aseries of claims which particularly point out and distinctly claim thesubject matter which Applicants consider to be their invention, acomplete understanding of the invention will be gained from thefollowing detailed description, which is given in connection with theaccompanying drawing, in which:

FIG. 1 is a partially schematic view of a typical high bypass, largediameter turbofan engine;

FIG. 2 is a cross-sectional view of a plurality of adjacent turbofanblades, in a given stage, taken along line 2-2 of FIG. 1;

FIG. 3 is a perspective view of a blade similar to those shown in FIG.2;

FIG. 4 is an enlarged partial, sectional view of a turbofan engineshowing an alternative noise reduction embodiment;

FIG. 5 is a partial cross-sectional view of a single blade taken alongline 5-5 of FIG. 4;

FIG. 6 is a partial, cross-sectional view of an alternate configurationof a turbofan blade; and 5 FIG. 7 is a graphical plot ofa noisesignature of a typical turbofan engine showing the effects of thepresent invention.

DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring now to the drawingwherein like numerals correspond to like elements throughout, referenceis directed initially to FIG. I wherein a gas turbine engine 10 of thehigh bypass, front fan type is shown to include a core engine 12, whichis essentially a turboshaft engine in that it includes a compressor 14,a combustor 16, a gas generator turbine I8 for driving the compressorl4, and a power turbine arranged in axially spaced serial flowrelationship. The inner turbomachine, or core engine 12, is enclosedwithin a cylindrical casing 22 which terminates at its downstream end inan exhaust nozzle 24 through which the combustion products of the coreengine 12 may be discharged to produce thrust. To provide additionalthrust, a fan 26 is mounted upstream of the core engine 12 and is drivenby the power turbine 20. The fan 26 is single stage comprised of aplurality of fan blades 28 which extend radially from a fan disc 30,which is coupled for rotation to the power turbine 20 by means ofa shaft31. The fan blades 28 extend radially across a bypass duct or passageway32 defined between an outer cylindrical casing 34 and a bullet nose" 36located upstream of the fan blades 28.

Downstream of the fan blades 28, the passageway 32 is divided intopassages 38 and 40 by a splitter platform 42. Radially positionedbetween the casing 34 and the splitter platform 42 are a plurality offan stator vanes 44, which are followed by a plurality of fan outletguide vanes 46. Thus, a portion of the air flow entering the passageway32 flows through the fan blades 28, into the passageway 38, through thestator vanes 44 and through the outlet guide vanes 46. This portion ofthe air flow thereafter exits through an outlet opening 48 formedbetween the casing 34 and the casing 22. Since this air is pressurizedin flowing through the fan blades 28, it provides additional forwardthrust to the turbofan engine 10.

The remainder of the air flowing through the passageway 32 and the fanblades 28 enters the passageway 40. As shown most clearly in FIG. 4,located within this passageway 40 are a plurality of inlet guide vanes50 for the core engine 12. Downstream of the inlet guide vanes 50 are aplurality of rotatable compressor blades 52, which extend from acompressor disc 54 and are coupled for rotation with the fan blades 28by means of the disc 54 and a shaft 56. Located downstream of thecompressor blades 42 is a row of stator vanes 58. Air passing throughthe stator vanes 58 flows into the core engine 12.

One characteristic of every high bypass ratio turbofan engine is thatthe diameter of the bypass fan is much larger than the diameter normallyassociated with a compressor or turbine section of the core engine.Because of this large diameter, the rotation of the fan 26 may result inhigh noise levels unless some provision is made for reducing theintensity of noise generated by such a fan. The present inventionrelates to a fan blade design which reduces the noise level associatedwith rotation of the large diameter fan. It will become obvious to thoseskilled in the art, however, that the presently disclosed concept can beapplied to any size blade for noise reduction purposes.

As shown in FIG. 2, the fan blades 28 include a cross section of airfoilshape having a leading edge 60, a trailing edge 62, a pressure orconcave surface 64 and a suction or convex surface 66. The fan blades 28are typical airfoil cross sections with the exception that a certainportion of material has been removed from the leading edges. That is,the leading edge 60 which would normally be associated with an airfoilsection having the pressure surface 64 and the suction surface 66 shownin FIG. 2 would extend as shown by the phantom lines 68. Said anotherway, the normal aerodynamic contour of the leading edge has beenflattened by removal of the material from the area designated by thenumeral 70 in FIG. 2.

The exact amount of blunting required would vary depending upon theparticular blade design and the application, but substantial bluntingwould appear to be the most promising from a noise reduction standpoint.In each case, a trade must be made between the efficiency of anaerodynamically ideal blade and the noise level of an ideally bluntedblade.

To enhance the reduction of noise generation of a bladed rotor stage byusing the present invention, it is preferred that each of the blades ofthe stage be treated as shown in FIG. 2. In this way, the blades of anyone stage (that is, carried by any one disc) will have their leadingedges disposed in a coplanar configuration, as "t line 69. Thisarrangement leads to a cumulative reduction in blade-noise generation bya given stage.

As shown in FIGS. 1 and 2, the blunting of the lead ing edge of the fanblades 28 may extend over the entire radial length of the blade 28. Inmany applications, however, it may be desirable to provide a blade 28',as shown in FIG. 3, wherein a first portion 72 of the leading edge isconstructed in normal fashion while a second portion 74 of the leadingedge is blunted. That is, the blunting would extend only over the outerportion of the leading edge, e.g., the outer one-third of the radialheight; while the remainder of the blade leading edge would beconstructed to have a normal aerodynamic contour. A blade constructed inaccordance with the design shown in FIG. 3 would minimize theperformance losses that might be associated with the blade constructedas shown in FIGS. I and 2 and would still effect a significant reductionin noise generation.

Instead of physically removing material from the leading edge of the fanblade 28 as described in connection with FIGS. I-J, the same noisereduction ef fects can be obtained by aerodynamically blunting theleading edge by injecting air into the leading edge of a bladeconstructed as shown in FIGS. 4 and 5. For this reason, a fan blade 28"is shown to include a chamber 76 which extends along the radial lengthof the blade 28" and is spaced a slight distance from the leading edgethereof (FIG. 5). A plurality of metering passages 78 are connected tothe chamber 76 at various radial heights along the blade 28".

As further shown in FIGS. 4 and 5, the metering passages 78 extend fromthe chamber 76 to the leading edge of the fan blade 28". The passage 78thus presents a path for the flow of fluid from the chamber 76 to thefan air stream through the leading edge of the fan blades 28". In orderto provide the chamber 76 with a suitable source of pressurized fluid,the fan wheel 30 is provided with a plurality of passageways 80 whichare adapted to align with the chamber 76 when the fan blades 28" arepositioned within suitable openings in the periphery of the fan wheel 30in a manner well known in the art.

The passageway 80 extends from the periphery of the disc 30 to a facethereof which lies in fluid flow communication with a chamber 82 formedby the disc 30, the disc 54, the shaft 56, and associated hardware asshown in FIG. 4. The chamber 82 is then pressurized with a suitablesource of high pressure air such as fan discharge or compressorinterstage bleed air. For this reason, tubing 84 extends from anysuitable source of pressurized air to the chamber 82. The tubing 84 isprovided with a valve 86 which permits selective pressurization of thechamber 82 for a purpose to be described.

As described above, blunting of the leading edge of the fan blades 28effectively reduces the noise generated by rotation of the fan. Asfurther described above, this blunting can be accomplished by removal ofmaterial from the leading edge of the fan blades 28. However, bluntingof the leading edge not only effectively reduces noise levels but canhave detrimental effects on performance of the fan. For this reason, itis desirable to provide a fan blade which includes a blunted leadingedge during those portions of flight in which noise is considered to bemost critical, i.e., take-off and landing approach, but which has anormal leading edge during the remaining portions of flight. The fanblade 28" shown in FIGS. 4 and 5 includes such capabilities. That is,during take-off and landing approach the valve 86 would be opened suchthat the chamber 82, and thus the chamber 76 within the fan blades,would be pressurized. In this manner, high pressure air would flowthrough the metering passageways 78 and out the leading edge of the fanblade 28". This efflux of pressurized air from the leading edge of thefan blades in a direction opposite that of the air flow within the fanduct 32 presents an effective blunted leading edge and causes noisereduction in much the same manner as does actual removal of materialfrom the fan blade leading edge.

It should be obvious to those skilled in the art that the outlet of themetering passage 78 could take many shapes and yet effectively provide ablunted leading edge. For example, the outlet could be formed as anelongated slot extending over some radial distance or could be formed ofa plurality of smaller circular holes at various spacings. Furthermore,as described in connection with FIG. 3, it may be necessary in manyapplications to provide metering passages only along the outer portionof the fan blades 28". This, of course, would reduce the complexity andcost associated with construction of such a blade.

It should further be obvious to those skilled in the art that the bluntleading edge could also be provided by many mechanical means. Forexample, a leading edge 68' of a fan blade 28" could be convered with adeformable member 88 as shown in FIG. 6. The member 88 would beconnected to the blade 28" in fluid sealing relationship at points 90and 92 and at the top and bottom of the blade. At certain radiallocations, the rigid material which forms the fan blade would be removedfrom the leading edge 68' to provide a surface 94 spaced a slightdistance from the deformable member 88. Thus, at certain radiallocations a small chamber 96 would exist between the deformable member88 and the surface 94.

The flexible member 88 would be shaped in the form of a normal leadingedge and would be made of material having sufficient rigidity tomaintain this shape during cruise operation. During take-off andlanding, however, the chamber 92 would be evacuated such that the member88 is drawn back against the surface 94. [n this manner, the member 88would provide a blunt leading edge during that portion of the flightwherein noise reduction is desirable while providing a clean leadingedge to eliminate detrimental performance effects during cruiseoperation. It should be obvious that similar effects could be gained bypressurizing the chamber behind the deformable member and thus expandingthe member to provide a blunt contour.

FIG. 7 is a graphical plot of actual test data showing the noisereduction associated with blunting a leading edge in a manner similar tothat shown in connection with FIGS. 1 and 2 above. Curve 1 is aone-third octave band plot of the second pressure level of an enginehaving fan blades with a normal leading edge. Curve ll is a similar plotof the sound pressure levels of an engine with a blunted leading edgefan blade wherein the transducer is located in the fan inlet. As can beseen from this graph, noise reduction of 5 to 12 SPL-dB were measured inthis one-third octave band above l,250HZ. These noise reductionsoccurred in both the inlet and the fan nozzle and occurred within boththe tone noise and the broad band noise.

While the actual mechanisms or aerodynamic phenomena which produce thenoise reduction as a result of blunting the leading edge are notcompletely understood, at least two theories attempting to define theseactual mechanisms can be hypotesized. At take-off power, the bluntedleading edge increases the strength of the bow shock thereby producing astate where noise generated behind the shock will have greaterdifficulties traveling forward. The stronger bow shock also results in alower passage (blade) Mach number which produces lower broad band noise.

At approach power settings, the pressure fields on the leading edge ofthe fan rotor are intensified by blunting at least the tip of the blade.This intensification produces a number of changes in local aerodynamics,such as redistribution of the flow and changes in local Mach number(pressure ratio and flow will decrease for a given fan rpm). Thisredistribution of flow evidently causes a reduction in broad band noiseassociated with the interaction of fan stream air and the leading edgeof the fan rotor.

What we claim is:

1. In a gas turbine engine of the type which includes rotatingturbomachinery, a combustion section, and an exhaust nozzle adapted toprovide a propulsive force, the improvement comprising:

a rotatable disc; and

a plurality of turbomachinery rotor blades carried by said disc andhaving leading edges at least a portion of which include blunting meansfor reducing the noise generated by the rotation of said blade, andwherein said leading edges of said blades are generally coplanar.

2. The improvement of claim 1 wherein said blunting means comprises aflattened leading edge which extends over at least a portion of thelength of each of said blades.

3. The improvement of claim 2 wherein said flattened leading edgeextends along approximately the outer one-third of the blade height.

4. The improved turbomachinery rotor blade recited in claim 1 whereinsaid blade comprises a fan blade of a turbofan engine.

5. The improvement of claim 2 wherein said flattened leading edge ismore blunt than that of an aerodynamically ideal blade in a selectedapplication.

1. In a gas turbine engine of the type which includes rotatingturbomachinery, a combustion section, and an exhaust nozzle adapted toprovide a propulsive force, the improvement comprising: a rotatabledisc; and a plurality of turbomachinery rotor blades carried by saiddisc and having leading edges at least a portion of which includeblunting means for reducing the noise generated by the rotation of saidblade, and wherein said leading edges of said blades are generallycoplanar.
 2. The improvement of claim 1 wherein said blunting meanscomprises a flattened leading edge which extends over at least a portionof the length of each of said blades.
 3. The improvement of claim 2wherein said flattened leading edge extends along approximately theouter one-third of the blade height.
 4. The improved turbomachineryrotor blade recited in claim 1 wherein said blade comprises a fan bladeof a turbofan engine.
 5. The improvement of claim 2 wherein saidflattened leading edge is more blunt than that of an aerodynamicallyideal blade in a selected application.